Conventionally, a rotorcraft rotor comprises a hub driven in rotation about an axis of rotation by a drive shaft or an outlet shaft from a main gearbox, together with at least three blades that are fastened to the hub via appropriate hinges, in particular via respective laminated spherical thrust bearings dedicated to each of the blades.
If each blade were to be fastened to a hub without being hinged, then the resulting rotor is a rigid rotor. With a rigid rotor, in hovering flight, the distribution of aerodynamic forces along a blade gives rise to a bending moment of very large magnitude at the root of the blade. In horizontal flight, the so-called “advancing” blade generates more lift than the so-called “retreating” blade because of the difference in air speeds, as explained in greater detail below.
Consequently, the resultant of the aerodynamic forces acting on a blade does not have the same value in all azimuth positions, nor does it have the same point of application: the bending moment at the root of the blade is thus large and varying, thereby generating alternating mechanical stresses, giving rise to a fatigue phenomenon that is harmful to materials. Furthermore, the resultant of the aerodynamic forces from all of the blades is no longer disposed along the drive axis of the rotor, thereby generating a roll moment, said roll moment increasing with increasing forward speed of the rotorcraft, which can make it difficult to keep the rotorcraft in equilibrium in horizontal flight.
In order to remedy those drawbacks, it is known to hinge each blade on the hub about a respective axis that is perpendicular to the drive shaft and referred to as the vertical flapping axis, which corresponds to a vertical flapping hinge that is capable of transferring lift but that can under no circumstances transfer a bending moment. Consequently, if a blade has a flapping hinge connection with the hub, then the flapping bending moment is zero at the point of attachment constituted by said flapping hinge. For the blade to be in equilibrium, centrifugal forces keep the blade extending outwards after it has risen a little so that the resultant of the lift and the centrifugal forces intersects said flapping axis, and allowing conicity a0 to appear.
Under such conditions, there is no longer a large roll moment in horizontal flight and the blades no longer rotate in a plane, but rather their outer ends describe a very flat cone. In practice, the flapping axis is then no longer on the axis of rotation but is rather offset therefrom by a distance a, referred to as eccentricity.
It should also be recalled that in order to provide a helicopter with lift in its various configurations, it is necessary to be able to control the lift provided by the rotor and to vary it. That is why a pitch hinge is also provided, about an axis that is substantially parallel to the span of the corresponding blade. This new degree of freedom enables the lift of the blade to be controlled by acting on the general pitch control, and also makes it possible to vary pitch cyclically, thereby enabling the plane of rotation of the blade to be controlled, which blades then describe a cone of an axis that no longer coincides with the drive axis: the resultant of the forces applied to the hub changes direction when the plane of the rotor changes. As a result, moments are applied about the center of gravity of the helicopter, thereby making it possible for it to be piloted.
As mentioned above, the plane of rotation of the blades may be other than a plane perpendicular to the drive shaft. Under such conditions, it is necessary for each blade to be hinged in drag since the end of each blade is at a distance from the rotor shaft that varies. Otherwise, inertial forces would generate alternating bending moments in the plane of each blade, thereby generating undesirable mechanical stresses. Such a drag hinge is provided by hinging a blade about a drag axis that is substantially parallel to the rotor axis, and consequently substantially perpendicular to drag forces. To enable such a blade to be driven by the drive shaft, it is naturally necessary for the drag hinge to be far enough away from the rotor axis to ensure that the moment due to centrifugal forces balances the moment due to drag and inertial forces, thus requiring the drag axis to be offset so as to present eccentricity e, and with it being necessary for the so-called “drag” angle δ not to be too large.
Consequently, the blades of a hinged rotor of a rotary wing aircraft, in particular of a helicopter, can move in the following four ways:
I) rotation about the rotor axis;
II) pivoting about a vertical flapping axis, made possible by the vertical flapping hinge;
III) pivoting about the drag axis, also referred to as the horizontal flapping axis, made possible by the horizontal flapping hinge or drag hinge; and
IV) pivoting about the axis of the blade made possible by the pitch hinge (with this not being specific to hinged rotors).
By way of example, as described in patent FR 2 497 073, the above pivoting movements II, III, and IV may be made possible by a single member such as a laminated spherical thrust-bearing.
Nevertheless, the oscillations of each blade about its drag axis may become coupled in unstable manner with the movements or the elastic deformation modes of the airframe, in particular the oscillations of a helicopter standing on the ground via its landing gear: this gives rise to a phenomenon known as “ground resonance” that can be dangerous for the aircraft when the resonant frequency of the oscillations of the blades about their drag axes is close to one of the resonant frequencies of oscillations of the aircraft on its landing gear.
Remedies to that phenomenon consist in introducing damping on the drag axes by means of a damper type device.
Such dampers include resilient return means of determined stiffness and damping qualities for combating resonance phenomena, in particular ground resonance, and also drive train resonance that can also appear, in particular in helicopters.
When rotor blades are excited in drag, the blades depart from their equilibrium positions and may become distributed unevenly in the circumferential direction, thereby creating an unbalance by moving the center of gravity of the rotor away from its axis of rotation. Furthermore, the blades that have moved away from their equilibrium positions oscillate about those positions at a frequency ωδ which is the resonant frequency of the blades in drag, and more exactly of the first mode of vibration in drag, referred to more simply as drag mode.
If Ω is the frequency of rotation of the rotor, it is known that the fuselage of the helicopter is thus excited at the frequencies |Ω±ωδ|.
When standing on the ground via its landing gear, the fuselage of the helicopter constitutes a system comprising a mass suspended above the ground by a spring and a damper constituted by the landing gear. The fuselage supported by its landing gear thus has its own resonant modes of vibration in roll and in pitching. There is a risk of instability on the ground when the resonant frequency of the fuselage on its landing gear, in roll or in pitching, is close to an excitation frequency, and in particular to the frequency |Ω−ωδ|, which corresponds to the ground resonance phenomenon. To avoid this instability, it is known firstly to seek to avoid these frequencies crossing before the rotor reaches its nominal speed of rotation, and if they cannot be prevented from crossing, then it is necessary to damp the movements of the fuselage on its landing gear sufficiently and also to damp the blades of the main rotor in their drag movements.
Consequently, the stiffness of the drag dampers of the blades of a main rotor must be selected so that the resulting resonant frequency of the blades in drag makes it possible to avoid a potential ground resonance zone, while simultaneously having sufficient damping. When the speed of rotation of the rotor passes through the critical speed, assuming said speed is lower than the nominal speed of rotation of the rotor, and regardless of whether the speed of rotation of the rotor is speeding up or slowing down, then the movements of the blades must be damped sufficiently to avoid entry into resonance.
That is why drag dampers with resilient return means of determined stiffness are also referred to as frequency adapters.
Document FR 2 672 947 describes a frequency adapter provided with a first cylinder that is elongate and blind, i.e. an elongate external strength member extending from a closed end to an open end that opens to the outside of the external strength member. A second elongate cylinder of the strength member type is then inserted through said open end into the inside of the first cylinder.
The frequency adapter then has resilient return means arranged between the first and second cylinders, specifically an elastomer ring that is bonded to the first and second cylinders.
Similarly, document FR 2 818 717 provides first, second, and third cylinders that are interleaved in one another, with two adjacent coaxial cylinders being bonded together by an elastomer ring.
More precisely, an elastomer ring connects the inside face of a side wall of the first cylinder to the outside face of the side wall of the second cylinder. Likewise, an elastomer ring connects the inside face of a side wall of the second cylinder to the outside face of the side wall of the third cylinder.
Documents EP 0 500 012 and WO 94/15113 also describe resilient return means arranged between two cylinders.
Consequently, the state of the art provides frequency adapters that are provided at least with first and second cylinders having respective elongate first and second side walls, the second cylinder being surrounded at least in part by the first cylinder, with resilient return means being bonded to the first and second side walls.
Such frequency adapters are very effective. Nevertheless, their resilient return means deteriorate over time. This deterioration is manifested by the appearance of cracks, fissures in the elastomers, thereby reducing the effectiveness of the adapter. Depending on the severity of the deterioration, it becomes necessary to replace the adapter.
The manufacturers of frequency adapters are consequently required to establish replacement criteria based on the dimensions of the cracks that can be seen from the outside in order to determine whether it has become necessary to replace a frequency adapter.
In order to monitor the physical integrity of such return means, an operator makes use of a small ruler. That method is made difficult to implement when accessibility to the return means is poor as a result of the large number of components in their vicinity, thereby considerably lengthening the time required for measurement and maintenance.
Document FR 2 860 582 relating to elastomer members provides for placing graduations on the visible surface of an elastomer. It would appear that this teaching applicable to elastomer members is incomplete in the context of a frequency adapter since the depth of the crack is not known.
The thickness of the return means of an adapter is considerable compared with its visible area, so transposing the teaching of document FR 2 860 582 to a frequency adapter would appear to be difficult.
Document U.S. Pat. No. 5,534,289 provides for placing two layers of colored microcapsules on a structure. Like document FR 2 860 582, the information provided by the method used remains fragmented, with it being difficult to evaluate the length and the depth of a crack.
Thus, the results of devices implemented in accordance with documents FR 2 860 582 and U.S. Pat. No. 5,534,289 may be difficult to interpret without some additional operation.
Document U.S. Pat. No. 5,493,899 provides for plunging an elastomer element in a solvent and observing whether the solvent penetrates into the inside of the element.
Finally, document U.S. Pat. No. 4,531,403 provides a method of detecting cracks in a solid material by measuring permeability.
The methods of documents U.S. Pat. No. 5,493,899 and U.S. Pat. No. 4,531,403 appear clearly to be ill-adapted to a frequency adapter arranged on a rotorcraft lift rotor.
The provisions set out in technical fields that are remote from the invention, i.e. remote from frequency adapters, do not, a priori, provide solutions that are completely satisfactory.
It should be observed that documents U.S. Pat. No. 5,205,710 and GB 1 568 455 relate to a blade and as a result they are remote from the technical field of the invention.
Furthermore, document DE 1 942 853 presents return means provided with two lateral grooves.